U.S. patent number 6,908,288 [Application Number 09/682,899] was granted by the patent office on 2005-06-21 for repair of advanced gas turbine blades.
This patent grant is currently assigned to General Electric Company. Invention is credited to Aaron Todd Frost, Shyh-Chin Huang, Melvin Robert Jackson, Charles Gitahi Mukira, Thomas Robert Raber, Raymond Alan White.
United States Patent |
6,908,288 |
Jackson , et al. |
June 21, 2005 |
Repair of advanced gas turbine blades
Abstract
Methods for repairing and manufacturing a gas turbine blade, and
the gas turbine blade repaired and manufactured with such methods
are presented with, for example, the repair method comprising
providing a gas turbine blade, the blade comprising a blade tip and
a blade body; removing at least one portion of the blade tip;
providing at least one freestanding tip insert; and disposing the
at least one tip insert onto the gas turbine blade body such that
the at least one tip insert replaces the at least one removed
portion of the blade tip.
Inventors: |
Jackson; Melvin Robert
(Niskayuna, NY), Frost; Aaron Todd (Lewisville, TX),
Huang; Shyh-Chin (Latham, NY), Mukira; Charles Gitahi
(Clifton Park, NY), Raber; Thomas Robert (Schenectady,
NY), White; Raymond Alan (Schenectady, NY) |
Assignee: |
General Electric Company
(Niskayuna, NY)
|
Family
ID: |
24741667 |
Appl.
No.: |
09/682,899 |
Filed: |
October 31, 2001 |
Current U.S.
Class: |
416/224;
29/889.1; 415/173.6; 415/200; 428/680 |
Current CPC
Class: |
F01D
5/005 (20130101); Y02T 50/67 (20130101); Y02T
50/672 (20130101); Y10T 428/12944 (20150115); Y10T
29/49318 (20150115); Y02T 50/60 (20130101) |
Current International
Class: |
F01D
5/00 (20060101); F01D 005/14 () |
Field of
Search: |
;416/224,241R,241A
;29/889.1-889.21,402.7-402.18,899.1,899.12,899.21,402.07
;415/173.6,200 ;428/680,937 ;148/428,400-405
;420/444-445,461-462 |
References Cited
[Referenced By]
U.S. Patent Documents
Other References
Rh-Base Refractory Superalloys for Ultra-High Temperature Use, Y.
Yamabe-Mitarai, Y. Koizumi, H. Murakami, Y. Ro, T. Maruko and H.
Harada, Scripta Materialia, vol. 36, No. 4, pp. 393-398, 1997.
.
Ir-Base Refractory Superalloys for Ultra-High Temperatures, Y.
Yamabe-Mitarai, Y. Ro, T. Maruko, and H. Harada, Metallurgical and
MAterials Transactions A, vol. 29A, Feb. 1998, pp.
537-549..
|
Primary Examiner: Look; Edward K.
Assistant Examiner: McAleenan; James M.
Attorney, Agent or Firm: DiConza; Paul J. Powell, III;
William E.
Claims
What is claimed is:
1. A method for repair of a gas turbine blade, comprising:
providing a gas turbine blade, said blade comprising a first
material and further comprising a blade tip and a blade body;
removing at least one portion of said blade tip; providing at least
one freestanding tip insert comprising a second material; and
disposing said at least one tip insert onto said gas turbine blade
body such that said at least one tip insert replaces said at least
one removed portion of said blade tip; wherein said second material
has at least one attribute selected from the group consisting of a.
a melting temperature greater than a melting temperature of said
first material by at least about 80.degree. C.; b. a fatigue life
at least about three times greater than a fatigue life of said
first material; and e. a creep life at least about three times
greater than that of said first material.
2. The method of claim 1, wherein said blade tip comprises at least
one squealer, and said at least one portion of said blade tip
comprises said at least one squealer.
3. The method of claim 1, wherein disposing comprises joining said
at least one tip insert to said blade by means of a process
selected from the group consisting of welding, brazing, and
diffusion bonding.
4. The method of claim 1, wherein said at least one tip insert
comprises at least one internal cooling channel.
5. The method of claim 1, wherein said at least one tip insert
comprises a plurality of cooling holes.
6. The method of claim 1, wherein said at least one tip insert
comprises a superalloy based on a metal selected from the group
consisting of iron, cobalt, and nickel.
7. The method of claim 6, wherein said at least one tip insert
comprises a directionally solidified material.
8. The method of claim 6, wherein said at least one tip insert
comprises a single crystal material.
9. The method of claim 1, wherein said blade comprises a first
material and said at least one tip insert comprises a second
material, and wherein each of a creep life, a fatigue life, and an
oxidation resistance for said first material is essentially
equivalent to each of a creep life, a fatigue life, and an
oxidation resistance of said second material, respectively.
10. The method of claim 1, wherein said second material comprises a
platinum group metal modified nickel-based superalloy.
11. The method of claim 10, wherein said superalloy comprises a
metal selected from the group consisting of Pt, Pd, Rh, Ir, and
Ru.
12. The method of claim 1, wherein said second material has an
oxidation resistance at least about 3 times greater than an
oxidation resistance of said first material.
13. The method of claim 12, wherein said second material comprises
a material selected from the group consisting of Rh, Pt, Pd, and
mixtures thereof.
14. The method of claim 13, wherein said at least one tip insert
further comprises a substrate material, and wherein said second
material is disposed on said substrate material.
15. The method of claim 14, wherein said second material comprises
a layer with a cross-sectional thickness in the range from about
0.13 mm to about 0.64 mm.
16. The method of claim 13, wherein said second material comprises
Rh at a level of at least about 65 atomic percent.
17. The method of claim 13, wherein said second material further
comprises a metal selected from the group consisting of Ir, Ru, and
mixtures thereof, at a level of up to about 5 atomic percent.
18. The method of claim 13, wherein said second material further
comprises Cr.
19. The method of claim 18, wherein the Cr is present at a level of
up to about 25 atomic percent.
20. The method of claim 18, wherein said second material further
comprises Al.
21. The method of claim 20, wherein said directionally solidified
eutectic material comprises Ni, Ta, and C.
22. The method of claim 20, wherein the Al is present at a level of
up to about 18 atomic percent.
23. The method of claim 20, wherein said second material further
comprises Ni.
24. The method of claim 23, wherein the Ni is present at a level of
up to about 45 atomic percent.
25. The method of claim 12, wherein said second material comprises
a refractory superalloy.
26. The method of claim 25, wherein said refractory superalloy
comprises Rh.
27. The method of claim 1, wherein said second material comprises a
directionally solidified eutectic material.
28. The method of claim 1, wherein said second material comprises
an oxide dispersion strengthened material.
29. The method of claim 28, wherein said oxide dispersion
strengthened material comprises Ni, Cr, and yttrium oxide.
30. A gas turbine blade repaired by the method of claim 1.
31. A method for repair of a gas turbine blade, comprising:
providing a gas turbine blade, said blade comprising a first
material and further comprising a blade tip and a blade body;
removing at least one portion of said blade tip; providing at least
one freestanding tip insert, said at least one tip insert
comprising a second material chosen from at least one of a single
crystal nickel-based superalloy, a NiTaC directionally solidified
cutcctic alloy, and an oxide dispersion strengthened alloy; wherein
said second material has at least one attribute selected from the
group consisting of a. fatigue life at least about three times
greater than a fatigue life of said first material, and b. a creep
life at least about three times greater than that of said first
material; and disposing said at least one tip insert onto said gas
turbine blade body such that said tip insert replaces said at least
one removed portion of said blade.
32. A method for repair of a gas turbine blade, comprising:
providing a gas turbine blade, said blade comprising a first
material and further comprising a blade tip said a blade body;
removing at least one portion of said blade tip; providing at least
one freestanding tip insert, said at least one tip insert
comprising a second material selected from the group consisting of
rhodium, platinum, palladium, and mixtures thereof, wherein said
second material has a melting temperature greater than a melting
temperature of said first material by at least about 80.degree. C.;
and disposing said at least one tip insert onto said gas turbine
blade body such that said tip insert replaces said at least one
removed portion of said blade.
33. A gas turbine blade comprising: a turbine blade body comprising
a first material; and a blade tip; wherein said blade tip comprises
at least one tip insert comprising a second material joined to said
blade body, and wherein said second material has at least one
attribute selected from the group consisting of a. a melting
temperature greater than a melting temperature of said first
material by at least about 80.degree. C.; b. a fatigue life at
least about times than a fatigue life of said first material; and
c. a creep life at least about three times greater than that of
said first material.
34. The gas turbine blade of claim 33, wherein a cross sectional
thickness of said at least one tip insert is less than a wall
thickness of said turbine blade body.
35. The gas turbine blade of claim 33, wherein a cross sectional
thickness of said at least one tip insert is at least equal to a
wall thickness of said turbine blade body.
36. The gas turbine blade of claim 33, wherein said at least one
blade tip comprises at least one squealer.
37. The gas turbine blade of claim 33, wherein said at least one
tip insert is joined to said blade body by means of a process
selected from the group consisting of welding, brazing, and
diffusion bonding.
38. The gas turbine blade of claim 33, wherein said at least one
tip insert comprises at least one internal cooling channel.
39. The gas turbine blade of claim 33, wherein said at least one
tip insert comprises a plurality of cooling holes.
40. The gas turbine blade of claim 33, wherein said at least one
tip insert comprises a superalloy based on a metal selected from
the group consisting of iron, cobalt, and nickel.
41. The gas turbine blade of claim 40, wherein said at least one
tip insert comprises a directionally solidified material.
42. The gas turbine blade of claim 40, wherein said at least one
tip insert comprises a single crystal material.
43. The gas turbine blade of claim 33, wherein said second material
comprises a platinum group metal modified nickel-based
superalloy.
44. The gas turbine blade of claim 43, wherein said superalloy
comprises a metal selected from the group consisting of Pt, Pd, Rh,
Ir, and Ru.
45. The gas turbine blade of claim 33, wherein said second material
has an oxidation resistance at least about three times greater than
an oxidation resistance of said first material.
46. The gas turbine blade of claim 45, wherein said second material
comprises a refractory superalloy.
47. The gas turbine blade of claim 46, wherein said refractory
superalloy comprises Rh.
48. The gas turbine blade of claim 33, wherein said at least one
tip insert further comprises a substrate material, and wherein said
second material is disposed on said substrate material.
49. The method of claim 48, wherein said second material comprises
a layer with a cross sectional thickness in the range from about
0.13 mm to about 0.64 mm.
50. The gas turbine blade of claim 33, wherein said second material
comprises Rh at a level of at least about 65 atomic percent.
51. The gas turbine blade of claim 33, wherein said second material
further comprises a metal selected from the group consisting of Ir,
Ru, and mixtures thereof, at a level of up to about 5 atomic
percent.
52. The gas turbine blade of claim 33, wherein said second material
further comprises Cr.
53. The gas turbine blade of claim 52, wherein the Cr is present at
a level of up to about 25 atomic percent.
54. The gas turbine blade of claim 52, wherein said second material
further comprises Al.
55. The gas turbine blade of claim 54, wherein the Al is present at
a level of up to about 18 atomic percent.
56. The gas turbine blade of claim 54, wherein said second material
further comprises Ni.
57. The gas turbine blade of claim 56, wherein the Ni is present at
a level of up to about 45 atomic percent.
58. The gas turbine blade of claim 33, wherein said directionally
solidified eutectic material comprises Ni, Ta, and C.
59. The gas turbine blade of claim 33, wherein said oxide
dispersion strengthened material comprises Ni, Cr, and yttrium
oxide.
60. A gas turbine blade comprising: a turbine blade body comprising
a first material; and a blade tip; wherein said blade tip comprises
at least one tip insert joined to said blade body, said at least
one tip insert comprising a second material chosen from at least
one of a single crystal nickel-based superalloy, a NiTaC
directionally solidified eutectic alloy, and an oxide dispersion
strengthened alloy, wherein said second material has at least one
attribute selected from the group consisting of a. a fatigue life
at least about three times greater than a fatigue life of said
first material, and b. a creep life at least about three times
greater than that of said first material.
61. A gas turbine blade comprising: a turbine blade body comprising
a first material; and a blade tip; wherein said blade tip comprises
at least one tip insert joined to said blade body, said at least
one tip insert comprising a second material selected from the group
consisting of rhodium, platinum, palladium, and mixtures thereof,
wherein said second material has a melting temperature greater than
a melting temperature of said first material by at least about
80.degree. C.
Description
BACKGROUND OF INVENTION
The present invention relates to components designed to operate at
high temperatures. More particularly, this invention relates to
methods for repair and manufacture of blades for gas turbine
engines, and the articles made and repaired from the use of these
methods.
In a gas turbine engine, compressed air is mixed with fuel in a
combustor and ignited, generating a flow of hot combustion gases
through one or more turbine stages that extract energy from the
gas, producing output power. Each turbine stage includes a stator
nozzle having vanes which direct the combustion gases against a
corresponding row of turbine blades extending radially outwardly
from a blade root, where a dovetail joint attaches the blade to a
supporting rotor disk, to a blade tip at the opposite end. The
blades are subject to substantial heat load, and, because the
efficiency of a gas turbine engine is proportional to gas
temperature, the continuous demand for efficiency improvements
translates to a demand for blades that are capable of withstanding
higher temperatures for longer service times.
Gas turbine blades are usually made of superalloys and are often
cooled by means of internal cooling chambers and the addition of
coatings, including thermal barrier coatings (TBC's) and
environmentally resistant coatings, to their external surfaces. The
term "superalloy" is usually intended to embrace iron-, cobalt-, or
nickel-based alloys, which include one or more other elements
including such non-limiting examples as aluminum, tungsten,
molybdenum, titanium, and iron. The internal air cooling of turbine
blades is often accomplished via a complex cooling scheme in which
cooling air flows through channels within the blade ("internal
cooling channels") and is then discharged through a configuration
of cooling holes at the blade surface. Convection cooling occurs
within the blade from heat transfer to the cooling air as it flows
through the internal air cooling channels. In more complex
configurations, fine internal orifices are often provided to direct
cooling air flow directly against inner surfaces of the blade to
achieve what is referred to as impingement cooling, while film
cooling is often accomplished at the blade surface by configuring
the cooling holes to discharge the cooling air flow across the
blade surface so that the surface is protected from direct contact
with the surrounding hot gases within the engine. TBC's comprise at
least a layer of thermally insulating ceramic and often include one
or more layers of metal-based, oxidation-resistant materials
("environmentally resistant coatings") underlying the insulating
ceramic for enhanced protection of the blade. Environmentally
resistant coatings are also frequently used without a TBC topcoat.
Technologies such as coatings and internal cooling have effectively
enhanced the performance of turbine blades, but material
degradation problems persist in turbine blades due to locally
aggressive conditions in areas such as blade tips.
A considerable amount of cooling air is often required to
sufficiently lower the surface temperature of a blade. However, the
casting process and the cores required to form the cooling channels
limit the complexity of the cooling scheme that can be formed
within a blade at the blade tip. The resulting restrictions in
cooling airflow often promote higher local temperatures in this
area relative to those existing in other locations on a given
blade. In typical jet engines, for example, bulk average blade
temperatures range between about 900.degree. C. to about
1000.degree. C., while blade tip surfaces often reach bout
1100.degree. C. or more. Maximum surface temperatures are expected
in future applications to be over about 1300.degree. C. Of
particular concern is the combination of stress with temperature,
because metals, including alloys used to make gas turbine blades,
tend to become weaker, or more easily deformed, as temperatures
increase. Thus, while stress of a certain level operating on a
cooler section of a blade may have little effect on performance,
the same stress level may be beyond the performance capability of
the material at hotter locations as described above. At such
elevated temperatures, materials are more susceptible to damage due
to a number of phenomena, including diffusion-controlled
deformation ("creep"), cyclic loading and unloading ("fatigue"),
chemical attack by the hot gas flow ("oxidation"), wear from
rubbing contact between blade tips and turbine shrouds, wear from
the impact of particles entrained in the gas flow ("erosion"), and
others.
Damage to blades, particularly at blade tips, leads to degradation
of turbine efficiency. As blade tips are deformed, oxidized, or
worn away, gaps between the blade tip and the turbine shroud become
excessively wide, allowing gas to leak through the turbine stages
without the flow of the gas being converted into mechanical energy.
When efficiency drops below specified levels, the turbine must be
removed from service for overhaul and refurbishment. A significant
portion of this refurbishment process is directed at the repair of
blade tips.
In current practice, the original blade tip material is made of the
same material as the rest of the original blade, often a superalloy
based on nickel or cobalt. Because this material was selected to
balance the design requirements of the entire blade, it is
generally not optimized to meet the special local requirements
demanded by conditions at the blade tip. The performance of alloys
commonly used for repair is comparable or inferior to that of the
material of the original component, depending upon the
microstructure, defect density, and chemistry of the repair
material. For example, many turbine blades are made using alloys
that have been directionally solidified. The directional
solidification process manipulates the orientation of metal
crystals, or grains, as the alloy is solidified from the molten
state, aligning the grains in one selected primary direction. The
resultant alloy has enhanced resistance to creep and fatigue during
service when compared to conventionally processed materials.
Advanced applications employ alloys made of a single crystal for
even further improvements in high-temperature creep and fatigue
behavior. However, when blade tips are repaired by some
conventional processes, using build-up of weld filler material, the
resulting microstructure of the repair is typical of welded
material, not directionally solidified or single-crystalline. Other
repair methods, such as applying powder mixtures wherein one powder
melts and densifies the repaired area during heat treatment,
results in microstructures that differ from that of the parent
alloy. Such microstructures, present in a conventional blade
material such as a superalloy, may cause the blade to require
excessively frequent repairs in advanced designs that rely on the
benefits of directional solidification or single crystal processing
to maintain performance.
Materials are characterized by several properties to aid in
determining their suitability for use in demanding applications
such as gas turbine blades. "Melting temperature" is used herein to
refer to the temperature at which liquid metal begins to form as
the material is heated. The term "creep life" is used in the art to
refer to the length of time until a standard specimen of material
extends to a preset length or fractures when subjected to a given
stress level at a given temperature. Similarly, the term "fatigue
life" is used in the art to describe the length of time until a
standard specimen fractures when subjected to a given set of
fatigue parameters, including minimum and maximum stress levels,
frequency of loading/unloading cycle, and others, at a given
temperature. The term "oxidation resistance" is used in the art to
refer to the amount of damage sustained by a material when exposed
to oxidizing environments, such as, for example, high temperature
gases containing oxygen. Oxidation resistance is generally measured
as the rate at which the weight of a specimen changes per unit
surface area during exposure at a given temperature. In many cases,
the weight change is measured to be a net loss in weight, as metal
is converted to oxide that later detaches and falls away from the
surface. In other cases, a specimen may gain weight if the oxide
tends to adhere to the specimen, or if the oxide forms within the
specimen, underneath the surface, a condition called "internal
oxidation." A material is said to have "higher" or "greater"
oxidation resistance than another if the material's rate of weight
change per unit surface area is closer to zero than that of the
other material for exposure to the same environment and
temperature. Numerically, oxidation resistance can be represented
by the time over which an oxidation test was run divided by the
absolute value of the weight change per unit area.
Materials particularly noted for high creep life include oxide
dispersion strengthened (ODS) materials and directionally
solidified eutectic (DSE) alloys. Several materials from these
classes have creep lives about three times those measured for
conventional superalloys. ODS materials use mechanical techniques
during processing to evenly distribute hard oxide particles of
sizes less than about 0.1 micron within a metallic matrix, with the
particles serving to make deformation of the material more
difficult. DSE alloys are characterized by carefully controlled
chemistry and processing, which produce a unique microstructure
comprising the inherent fibrous or, in some cases, lamellar
structure of the eutectic phase, with the fibers or lamellae
aligned along a desired axis of the cast part in a manner analogous
to a fiber-reinforced composite. DSE materials are also notable for
excellent fatigue life, with certain alloys having about three
times the fatigue lives measured for conventional superalloys. The
careful processing controls needed to produce ODS and DSE alloys
cause these materials to be prohibitively expensive.
The so-called "platinum group" of metal elements comprises rhodium
(Rh), osmium (Os), platinum (Pt), iridium (Ir), ruthenium (Ru),
palladium (Pd), and rhenium (Re)elements noted for high chemical
resistance and very high melting temperatures in comparison to
conventional superalloys. Several elements from this group are
noteworthy as examples of materials with substantially higher
oxidation resistance relative to current blade materials. Some
platinum group metals and several alloys based on platinum group
metals possess excellent resistance to oxidation at temperatures
exceeding the capabilities of many Ni-based superalloys. The class
of materials referred to as "refractory superalloys" offer
additional strength over the platinum group metals, though at the
expense of some oxidation resistance. These alloys are based on Ir
or Rh, with transition metal additions of up to about 20 atomic
percent, and are strengthened by a precipitate phase of generic
formula M.sub.3 X, where M is Rh or Ir and X is typically Ti, V,
Ta, or Zr, or combinations thereof. Some alloys of this type can
withstand 1-2 hour exposures to at least about 1600.degree. C.
without catastrophic oxidation. Creep life and fatigue life data
for these alloys are not readily available currently, but the high
strength of these alloys suggests they are superior to some degree
over conventional superalloys in both creep life and fatigue life
at the temperatures and stress levels relevant to gas turbine blade
components, although not to the same degree as the best ODS and DSE
alloys.
Platinum group metals also have been incorporated into conventional
superalloy compositions to produce a class of alloys, herein
referred to as "platinum-group metal modified superalloys", having
enhanced oxidation resistance and comparable mechanical properties
to conventional superalloys. Typical alloys of this class comprise
a conventional superalloy composition to which is added up to about
7 atomic percent of a platinum group metal, such as Ir, Rh, Pt, Pd,
and Ru. Use of materials incorporating platinum-group metals has
been limited to date due to the high density and very high cost of
these materials in comparison to more conventional blade
materials.
SUMMARY OF INVENTION
The selection of a particular alloy for use in a given turbine
blade design is accomplished based on the critical design
requirements for a number of material properties, including
strength, toughness, environmental resistance, weight, cost, and
others. When one alloy is used to construct the entire blade,
compromises must be made in the performance of the blade because no
single alloy possesses ideal values for the long list of properties
required for the application, and because conditions of
temperature, stress, impingement of foreign matter, and other
factors are not uniform over the entire blade surface.
It would be advantageous if the performance of both newly
manufactured and repaired blades could be improved to better
withstand the localized aggressive stress-temperature combinations
present at blade tips. However, it would not be desirable if
improvements to such properties as creep life, fatigue life, and
oxidation resistance were effected at the expense of other design
critical requirements of the turbine blade. For example, a blade
made entirely of platinum would have excellent oxidation
resistance, but would lack needed strength and would cost many
times the price of a blade made of conventional superalloy
material. Therefore, it would be beneficial if turbine blades could
be improved in a manner that would allow for enhanced blade tip
performance without significantly detracting from the overall
performance of the turbine blade. Furthermore, it would be
advantageous if methods of turbine blade repair and manufacture
could be developed that would overcome the limitations of typical
weld repair and powder deposition methods described above by
allowing for new blade tips to comprise material with properties
equal to, and often greater than, those of the original blade.
The present invention provides several embodiments that address
this need for blades with improved performance. One embodiment
provides a method for repair of a gas turbine blade, the method
comprising providing a gas turbine blade, the blade comprising a
blade tip and a blade body; removing at least one portion of the
blade tip; providing at least one freestanding tip insert; and
disposing the at least one tip insert onto the gas turbine blade
body to form a new blade tip. A second embodiment provides a method
for manufacturing a gas turbine blade, the method comprising
providing a gas turbine blade body, providing at least one
freestanding tip insert, and disposing the at least one tip insert
onto the gas turbine blade body such that a blade tip of the
turbine blade comprises the at least one tip insert. A third
embodiment provides a freestanding tip insert for manufacture and
repair of a tip of a gas turbine blade, the tip insert comprising
an external surface that substantially conforms with specified
nominal dimensions for an external surface of the blade. A fourth
embodiment provides a gas turbine blade comprising a turbine blade
body and a blade tip, wherein the blade tip comprises at least one
freestanding tip insert affixed to the blade body.
BRIEF DESCRIPTION OF DRAWINGS
These and other features, aspects, and advantages of the present
invention will become better understood when the following detailed
description is read with reference to the accompanying drawings in
which like characters represent like parts throughout the drawings,
wherein:
FIG. 1 is a perspective view of a gas turbine blade equipped with a
blade tip in accordance with embodiments of the present
invention;
FIG. 2 is a cross-sectional view of a freestanding tip insert
suitable for use as the blade tip of FIG. 1;
FIG. 3 is a cross-sectional view of one possible configuration for
the joint between the blade tip and the blade body in accordance
with embodiments of the present invention;
FIG. 4 is a cross-sectional view of another possible configuration
for the joint of FIG. 3;
FIG. 5 is a cross-sectional view of an example of an airfoil tip
with squealers;
FIG. 6 is a cross-sectional view of an example of a tip insert;
FIG. 7 is a cross-sectional view of another example of a tip
insert; and
FIG. 8 is a graph of oxidation data.
DETAILED DESCRIPTION
Structure, manufacture, and repair embodiments of the present
invention are useful for components that operate at elevated
temperatures, and particularly turbine blades (also referred to as
"buckets") for gas turbine engines wherein the maximum metal
temperatures typically range from about 1000.degree. C. to over
about 1200.degree. C. in current systems and temperatures over
about 1300.degree. C. are envisioned for future applications.
Referring to FIGS. 1 and 2, one embodiment of the invention
provides a method for repair of a gas turbine blade 10, the method
comprising: providing a gas turbine blade 10, the blade comprising
a blade tip 11 and a blade body 12; removing at least one portion
of the blade tip 11; providing at least one freestanding tip insert
20 (FIG. 2); and disposing the at least one tip insert 20 onto the
gas turbine blade body 12 such that the at least one tip insert 20
replaces the at least one portion of the blade tip 11. The term
"tip insert" used herein refers to a freestanding article suitable
in size, shape, and material properties to be used as a blade tip
11 or a portion of a blade tip 11 upon being disposed onto a blade
body 12. The term "replaces" used herein means the tip insert 20
occupies the position on the blade 10 formerly occupied by the at
least one removed portion. The height 13 of the blade tip 11 is in
the range from about 3 mm to about 4 mm for blades used in aircraft
engines, and from about 4 mm to about 11 mm for blades used in
land-based power generation turbines. Several possible
configurations exist for the joint between the at least one tip
insert 20 and the blade body 12 (FIG. 1), two of which are
illustrated in FIGS. 3 and 4. Those skilled in the art will
appreciate from FIG. 3 that in certain embodiments the cross
sectional thickness 21 of the at least one tip insert 20 is less
than the wall thickness 30 of the blade body 12, while in other
embodiments, as depicted in FIG. 4, for example, the cross
sectional thickness 21 of the at least one tip insert 20 is at
least equal to the wall thickness of 30 of the blade body 12. In
particular embodiments, a non-limiting example of which is shown in
FIG. 5, the blade tip 11 comprises at least one "squealer" 50 (a
protrusion of material from the tip cap 51), and the at least one
portion 52 of said blade tip 11 removed and replaced in the method
of the present invention comprises said at least one squealer 50.
Squealers 50 are employed to enhance the seal between the rotating
turbine blade 10 and the adjacent stator (not shown), and because
they are often subject to rubbing against the shroud during
operation, squealers 50 undergo heating from both the hot gas
stream and frictional heating from rubbing contact. Therefore, a
squealer 50 is likely to be a section of the blade tip 11 highly
prone to damage and highly likely to need replacement in accordance
with the present invention. The embodiment depicted in FIG. 5 shows
a blade tip 11 with two squealers 50, each of which having a
portion 52 that has been removed and then replaced in accordance
with the present invention.
In certain embodiments, the disposing step comprises joining the at
least one tip insert 20 to the blade 10 by means of a process
selected from the group consisting of welding, brazing, and
diffusion bonding. The joint line 31 between the at least one tip
insert 20 and the blade body 12 is advantageously located in a
position sufficiently close to a cooling channel 32 so that the
properties of the material along the joint line 31 are adequate to
perform under the stress temperature combinations everywhere along
the joint line 31. Alternatively, the height 13 (FIG. 1) of the at
least one tip insert is sufficiently large to place the joint line
31 in a region on the blade 10 where the stress-temperature
combination is able to be accommodated by the material along the
joint line 31.
The dimensions of the blade 10 (FIG. 1) depend upon the particular
blade design under consideration. Typically, design parameters such
as, for example, the blade tip height 13 and the shape of an
airfoil external surface 34 (FIG. 3), have preferred, or "nominal"
values and tolerance ranges documented in technical specifications
for the purposes of quality control. Thus it is commonly accepted
in the art to refer to, for example, "specified nominal dimensions"
for an external surface 34 of an airfoil, which in this example
would signify the nominal shape documented for the airfoil external
surface 34 in the pertinent technical specification.
In some cases, special consideration regarding the selection of the
joining process is required, especially where large differences in
melting point exist between material comprising the blade body 12
and the material comprising the at least one tip insert 20. For
example, where the blade body material comprises a Ni-based
superalloy with an exemplary melting temperature range from about
1300.degree. C. to about 1350.degree. C. and the tip insert
material comprises a large amount of a platinum-group metal, with a
melting temperature of at least about 1500.degree. C., the
employment of a joining process that generates low heat input into
the blade body is useful to avoid overheating the blade body
material. For example, brazing techniques and diffusion bonding
processes are less likely to overheat the turbine blade than
welding processes.
Diffusion bonding comprises bringing the components to be joined
into intimate contact and heating them to a sufficiently high
temperature such that solid-state diffusion occurs at the interface
between the two components, forming a continuous solid bond.
Fixtures (not shown) are used to ensure intimate contact is
maintained throughout the procedure. In one embodiment of the
present invention, the temperature is at least about 1200.degree.
C. so that a suitable bond can be achieved in about 4 hours.
In certain embodiments, after the at least one tip insert 20 has
been disposed onto the blade body, further process steps are used
to ensure the repaired blade meets design requirements for proper
operation. Examples of such requirements include, but are not
limited to, surface finish specifications, dimensional
requirements, and bond strength requirements for the bond joining
the tip insert material to the blade body 12. In one embodiment, a
step of heat treating the repaired blade 10 is used to improve the
bonding between the at least one tip insert 20 and the blade body
12, to relieve stresses accumulated by the repair process, and to
improve the metallurgical condition of the overall part in terms of
its grain size and precipitate phase distribution. Such a heat
treatment step is typically done in vacuum or in an inert gas to
avoid oxidizing the part, and is carried out using the process
specified for the particular alloy comprising the blade body
material to ensure the alloy's metallurgical properties are within
the range specified for the turbine blade when processing is
completed. In other embodiments, the step of disposing the at least
one tip insert 20 comprises one or more machining operations,
including grinding, milling, or other such processes, to restore
the blade 10 to specified final dimensions and surface finish
requirements. In certain embodiments this machining step includes a
process such as grinding to provide a surface finish for the
repaired turbine blade that meets the pertinent specification
limit. Particular embodiments have a coating 23 (FIG. 2) applied to
the turbine blade to afford even further high-temperature
protection. This coating 23 typically comprises at least one layer,
and optionally, in the case where a combination of a ceramic
thermal barrier coating and an environmentally resistant coating is
employed, a plurality of layers.
The tip insert 20 used in certain embodiments of this invention
further comprises at least one internal cooling channel 32, and in
particular embodiments the tip insert further comprises a plurality
of cooling holes 24 (FIG. 2). The cooling holes 24 are created
using any one of a number of techniques, including, for example,
laser drilling, electric discharge machining, and electron beam
drilling. These features add to the ability of the finished blade
to perform under demanding thermal conditions.
In specific embodiments, the at least one tip insert 20 comprises a
superalloy based on (i.e., the single largest elemental component
by weight) a metal selected from the group consisting of cobalt,
iron, and nickel. The at least one tip insert comprises a
directionally solidified material in particular embodiments, and in
selected embodiments the at least one tip insert comprises a single
crystal material. The directionally solidified and single crystal
embodiments are provided to enhance the high temperature
performance of the blade tip during service.
The blade tip 11 often reaches temperatures over about 200.degree.
C. higher than the average temperature of the blade 10. Because of
the particularly aggressive combination of stress and temperature
present at blade tips, primary performance characteristics required
for materials in these sections include creep life, fatigue life,
oxidation resistance, and melting temperature, for example. The
temperature range of interest in references to creep life, fatigue
life, and oxidation resistance herein includes the range of from
about 900.degree. C. to about 1200.degree. C., and relative
statements made herein comparing creep rupture lives, fatigue
lives, oxidation resistances, and melting temperatures of various
materials assume equivalent levels of stress, temperature, and
other critical factors for each material being compared.
In certain embodiments, the blade 10 comprises a first material and
the at least one tip insert 20 comprises a second material, and
each of a creep life, a fatigue life, and an oxidation resistance
for the first material is essentially equivalent to each of a creep
life, a fatigue life, and an oxidation resistance of the second
material, respectively. The term "essentially equivalent" used
herein means within the interval from about 20% below the value for
the first material to about 20% above the value for the first
material. Having these properties for the first and second
materials be essentially equivalent in a repair method represents
an improvement over conventional repair methods described above, in
that the original performance levels for the blade are
restored.
In particular embodiments, the blade comprises a first material and
the tip insert comprises a second material, and at least one
material property for the second material exceeds a corresponding
material property for the first material, the at least one material
property selected from the group consisting of oxidation
resistance, creep life, and fatigue life. The term "exceeds" as
used herein means that the pertinent material property for the
second material has a value that is at least about 120% of the
value of the corresponding property of the first material. Specific
embodiments of this type include cases where the second material
comprises a platinum group metal modified nickel-based superalloy.
These embodiments include particular cases where the platinum group
metal modified nickel-based superalloy comprises a metal selected
from the group consisting of Pt, Pd, Rh, Ir, and Ru. A non-limiting
example of an alloy of this type suitable for use in embodiments of
the present invention comprises the following nominal composition
(in weight percent): about 5.1% aluminum, about 8.0% chromium,
about 11.4% tungsten, about 1.7% titanium, about 1.5% niobium,
about 9.5% cobalt, about 0.09% carbon, about 0.02% boron, about
0.07% zirconium, about 9.9% platinum, and the balance nickel.
Other embodiments are provided by the present invention in which
the second material has an oxidation resistance at least about
three times greater (i.e., a measured weight change rate that is at
most 33% of that measured for the first material under the same
conditions) than an oxidation resistance of the first material.
Materials selected from the group consisting of rhodium (Rh),
platinum (Pt), palladium (Pd), and mixtures thereof show sufficient
oxidation resistance for use as the second material in these
embodiments. In particular embodiments of the present invention,
the second material comprises Rh at a level of at least about 65
atomic percent.
In selected embodiments, non-limiting examples of which are shown
in FIGS. 6 and 7, the at least one tip insert 60 further comprises
a substrate material 62, and the second material 64 is disposed on
the substrate material 62. In particular embodiments, the second
material 64 comprises a layer with a cross-sectional thickness 72
(FIG. 7) in the range from about 0.13 mm to about 0.64 mm. The
substrate material 62 comprises a material with strength
properties, such as, for example, creep life and fatigue life, that
are higher than those of the second material 64. These embodiments
are provided for applications where the stress level present at the
blade tip during service is too high to be fully supported by the
platinum group metals listed above. The second material 64 provides
superior oxidation resistance while allowing the stronger substrate
material 62 to take up most of the load.
FIG. 8 is a graph showing oxidation data for Pt, Rh, Pd, and
iridium (Ir) as well as for two conventional turbine blade
materials, a directionally solidified Ni-base superalloy and a
single crystal Ni-base superalloy of a different composition than
the former alloy. The test performed to generate this data used
metal specimens that were identical in size and each was exposed
for the same amount of time (1000 hours), and so the change in
specimen diameter is plotted as a direct measure of oxidation
effects in order to compare material performance. Although each of
the platinum group metals showed lower losses of metal than would
be expected for the two superalloys for temperatures above about
1300.degree. C., Ir was the worst of the platinum group metals
tested and its oxidation rate was deemed to be too high for use as
a major (>30 atomic percent) component of the second material
described for the present invention. Ruthenium (Ru) shows similar
performance to Ir under these conditions. However, these elements
are useful as minor alloying additions, and in certain embodiments
of the invention, the second material comprises a metal selected
from the group consisting of Ir, Ru, and mixtures thereof, at a
level of up to about 5 atomic percent.
The second material, as employed in some embodiments of the present
invention, further comprises chromium (Cr), which provides
additional oxidation resistance to the material. In particular
embodiments, the Cr is present at a level of up to about 25 atomic
percent. Certain embodiments further comprise aluminum (Al), and in
particular embodiments the Al is present at a level of up to about
18 atomic percent. In certain embodiments, the second material
further comprises nickel (Ni), which in certain embodiments is
present at a level of up to about 45 atomic percent.
Still other embodiments employ a second material comprising a
refractory superalloy, and in particular embodiments, the
refractory superalloy comprises rhodium (Rh). In particular
embodiments of the present invention, the second material comprises
Rh at a level of at least about 65 atomic percent. Alloys
comprising iridium at such levels have been shown to exhibit poor
oxidation resistance in the environments under consideration for
the present invention, and thus in particular embodiments of the
invention, the refractory superalloy comprises Ir at a level of at
most about 5 atomic percent. To attain the required mechanical
properties, the refractory superalloys further comprise a quantity
of additional material, where the quantity of additional material
comprises at least one supplementary element selected from the
group consisting of titanium (Ti), vanadium (V), zirconium (Zr),
niobium (Nb), molybdenum (Mo), hafnium (Hf), tantalum (Ta),
tungsten (W), and mixtures thereof. These alloys do not resist
oxidation to the levels shown by the Pt-group metals themselves,
but their combination of oxidation resistance and mechanical
properties is often suitable for use in embodiments of the present
invention. In particular embodiments, the at least one supplemental
element is present in the second material at a level of up to about
7 atomic percent. In certain embodiments, the quantity of
additional material comprises a plurality of supplemental elements,
each supplemental element selected from the same group as listed
above for the at least one supplementary element. In certain of
these embodiments comprising a plurality of supplemental elements,
the quantity of additional material is present in the second
material at a level at or below about 10 atomic percent.
In certain embodiments of the present invention, the creep life of
the second material is greater than the creep life of the first
material. In other embodiments, the fatigue life of the second
material is greater than the fatigue life of said first material.
In particular embodiments, the creep life of the second material is
at least about three times greater than the creep life of the first
material, and in certain embodiments the fatigue life of the second
material is at least about three times greater than the fatigue
life of the first material. With a material having higher creep
life or higher fatigue life in place at the blade tip, a turbine
blade is better able to withstand the severe environment existing
locally at these locations than even a new conventional turbine
blade.
The second material comprises a directionally solidified eutectic
(DSE) in certain embodiments of the invention. Particular
embodiments of this type employ a DSE comprising nickel (Ni),
tantalum (Ta), and carbon (C), herein referred to as "NiTaC", and
an exemplary NiTaC composition is shown in the Table. NiTaC alloys
of this type form a fibrous microstructure with very strong and
hard tantalum carbide fiber-shaped phase reinforcing a more ductile
Ni-based metallic matrix phase. The table also displays an
alternate DSE composition, referred to herein as NiNbC, where
niobium (Nb) is used as the carbide forming element in place of Ta.
DSE Alloys of the types exemplified by the compositions displayed
in the Table exhibit creep rupture lives exceeding those of
commonly used single-crystal superalloys by a factor in the range
from about 2 to about 10, where the test load is about 21 MPa at a
temperature of about 1150.degree. C. Fatigue lives for these
exemplary alloys exceed those of commonly used single crystal
alloys by a factor in the range of from about 1.5 to about 5 at a
temperature of about 1150.degree. C., where the strain range is
about 0.1% and the frequency is about 20 cycles per minute.
[t1]
[t1] Alloy (w/o) Ni Cr Co Al W Mo Re Ta Nb C B Zr Ti Fe Y2O3 NiTaC
Bal 4.2 3.9 5.5 4.5 3.2 6.8 9.0 0 27 .01 0 0 0 0 NiNbC Bal 4 10 6
10 0 0 0 3.8 .5 0 0 0 0 0 MA754 Bal 20 0 .3 0 0 0 0 0 .05 0 0 .5 1
.6 MA6000 Bal 15 0 4.5 4 2 0 2 0 .05 .01 .15 2.5 0 1.1
In some embodiments where the second material creep life is higher
than that of the first material, the second material comprises an
oxide dispersion strengthened (ODS) material. Two exemplary
compositions are displayed in the Table. The alloy names MA754 and
MA6000 are trademarks of Inco, Limited. In particular embodiments
of this type, the ODS material comprises Ni, chromium (Cr),
aluminum (Al), and yttium oxide. The uniform dispersion of
sub-micron-sized yttrium oxide particles, typically present in a
concentration range of from about 0.5 volume % to about 2.5 volume
%, and their large-grained and elongated-grained microstructures,
provide a remarkably stable and effective barrier to dislocation
motion, accounting for the excellent creep life of these materials.
ODS alloys of the types exemplified by the compositions displayed
in the Table exhibit creep rupture lives exceeding those of
commonly used single-crystal superalloys by a factor in the range
from about 2 to about 10, where the test load is about 21 MPa at a
temperature of about 1150.degree. C. The chromium in the alloys,
present from about 15 weight % to about 20 weight %, provides
effective oxidation resistance to the Ni-based matrix.
The collection of alloys discussed above does not represent an
exhaustive list of all possible materials that may be employed to
form embodiments of the present invention. These materials are
discussed in order to illustrate the concepts of the present
invention and the manner in which their properties can be
advantageously exploited to achieve improved turbine blade life. A
significant benefit of embodiments of the present invention is that
the advantages of the second material are applied without
sacrificing the overall design requirements of the blade, because
the second material is disposed only at the blade tip, while the
remainder of the turbine blade comprises first material, selected
in accordance with the pertinent design requirements for the
particular turbine blade. Potentially disadvantageous properties of
certain second materials, such as high cost or density, have a
reduced effect on the overall blade because the second material
comprises only a fraction of the overall surface area of the blade.
The properties of the blade are thus "tailored" to the expected
localized environments, reducing the need for compromise during the
design process for new blades, and increasing the expected lifetime
for repaired articles operating in current systems.
Another embodiment of the invention provides a method for
manufacturing a gas turbine blade 10, the method comprising
providing a gas turbine blade body 12, providing at least one
freestanding tip insert 20 (FIG. 2), and disposing the at least one
tip insert 20 onto the gas turbine blade body 12 such that a blade
tip 11 of the turbine blade 10 comprises the at least one tip
insert 20. As in previously described embodiments, manufacturing
method embodiments are provided in which the blade body 12
comprises a first material and the at least one tip insert 20
comprises a second material, and the first and second materials
have the same characteristics as the first and second materials in
the previously discussed embodiments. The alternatives for the
composition of the second material discussed previously for the
repair method embodiments are also applied in certain of the
manufacturing method embodiments. Furthermore, the alternative
embodiments relating to aspects of the disposing step, set forth
above for the repair method embodiments, are also pertinent to the
manufacturing method embodiments of the present invention. In
certain embodiments, heat treatment and machining steps are
included in the manufacturing process, for the same reasons as
described above for repair method embodiments, as is the step of
applying a coating to the turbine blade to afford further
high-temperature protection. Particular embodiments provide that
the at least one tip insert 20 further comprises at least one
internal channel 32 (FIG. 2) to allow for internal air cooling of
the blade tip during service, and as described above for the repair
method embodiments, certain embodiments of the invention include a
plurality of cooling holes 24 (FIG. 2) in the at least one tip
insert 20.
Embodiments of the present invention also provide a freestanding
tip insert 60 (FIG. 6) for manufacture and repair of a tip 11 of a
gas turbine blade 10 (FIG. 1), the tip insert 60 comprising an
external surface 66 (FIG. 6) that substantially conforms with
specified nominal dimensions for an external surface 34 (FIG. 3) of
the blade 10. An "external surface" as used herein in reference to
the tip insert 60 refers to any surface of the tip insert 60 that,
when the tip insert is disposed on the turbine blade 10 in
accordance with embodiments of the present invention, is in contact
with the hot gas stream during operation of the gas turbine. An
"external surface" as used herein in reference to a blade 10 refers
to any surface that is in contact with the hot gas stream during
operation of the gas turbine. The term "substantially conforms" as
used herein means that upon disposition of the tip insert 60 onto
the blade 10, the resulting external surface 34 is within specified
tolerance limits for the blade at the blade tip, either as-disposed
or with routine machining to blend the surfaces.
In certain embodiments, the tip insert 20 comprises a material
having a creep life of at least about 1000 hours tested at about
1150.degree. C. and about 21 MPa, a fatigue life of at least about
33,000 cycles to failure tested at about 20 cycles per minute and a
strain range of about 0.1% at about 1150.degree. C., and an
oxidation resistance of at least about 6 h-cm2/mg at about
1150.degree. C. In specific embodiments, the tip insert material
has an oxidation resistance of at least about 20 h-cm2/mg at about
1150.degree. C., in others the tip insert material has a creep life
of at least about 3000 hours tested at about 1150.degree. C. and
about 21 MPa, and in others the tip insert material has a fatigue
life of at least about 100,000 cycles to failure tested at about 20
cycles per minute and a strain range of about 0.1% at about
1150.degree. C. Suitable alternatives for the composition of the
tip insert material, discussed previously for the "second" material
in the repair method and manufacturing method embodiments, are also
applied in certain tip insert embodiments. Certain tip insert
embodiments provide that the tip insert 20 further comprises at
least one cooling channel 32, and other embodiments provide that
the tip insert 20 further comprises a plurality of cooling holes 24
as described for the above embodiments.
According to the embodiments of this invention, any of a variety of
metal fabrication and processing methods is suitable to fabricate
the tip insert. Examples of suitable processes include, but are not
limited to, casting (including directional solidification and
single crystal processing methods); forging; extruding; in-situ
processing of braze tape; or forming on a sacrificial mandrel by
deposition processes such as electron beam physical vapor
deposition, laser powder consolidation, chemical vapor deposition,
ion plasma deposition, thermal spraying, and electroplating.
A gas turbine blade 10 embodiment of the present invention
comprises a turbine blade body 12 and a blade tip 11, wherein the
blade tip 11 comprises at least one tip insert 20 joined to the
blade body 12. As above, embodiments are provided in which the
blade body 12 comprises a first material, and the at least one tip
insert 20 comprises a second material, respectively, and these
materials have the characteristics described for the first material
and the second material, respectively, in previously discussed
embodiments. The alternatives for the composition of the tip insert
20 material, discussed previously for the repair method,
manufacturing method, and tip insert embodiments, are also applied
in certain turbine blade embodiments. Embodiments relating to the
existence of at least one cooling channel 32 in the insert 20 and a
plurality of cooling holes 24, and coating 23, as described for the
above embodiments, are also provided for certain turbine blade
embodiments.
While only certain features of the invention have been illustrated
and described herein, many modifications and changes will occur to
those skilled in the art. It is, therefore, to be understood that
the appended claims are intended to cover all such modifications
and changes as fall within the true spirit of the invention.
* * * * *